In launch and space vehicle applications, such as satellite launchers, satellites and other spacecrafts, liquid propellant thrusters, liquid propellant rocket engines and liquid propellant gas generators are often used. Such thrusters and rocket engines can for example be used for the purpose of orbit manoeuvring and attitude control, of satellites, and for example roll control and propellant settling in the main propulsion system of other space vehicles, in which case the rocket engines, or thrusters are often used in continues firing, off-modulation firing, pulse mode and single pulse firing, the duration of which typically can be fractions of a second to an hour. For such purposes small rocket engines, or thrusters are commonly used with a thrust of typically from 0.5 N to about 1.5 kN.
Such thrusters may be operated on ammonium dinitramide-(ADN)-based, liquid monopropellants, such as described in WO 2002/096832, and in WO 2012/166046. Some of the ADN-based, liquid monopropellants, are also being referred to as High Performance Green Propulsion (HPGP) monopropellants.
A reactor for the above ADN-based, liquid monopropellants has been described in WO 02/095207, as well as a thruster comprising the reactor. Such thrusters are also being referred to as HPGP thrusters.
The present inventors have found that, when operating a thruster as described in WO 02/095207 in certain pulse modes, hard starts are encountered. A hard start implies an overpressure condition during the ignition of the propellant in the thruster. In the worst cases, this takes the form of an explosion. A single hard start is obviously detrimental to the engine, and in worst case even fatal. The problem of hard starts has been observed for thrusters of 5N, 22 N and 200 N when operating the thruster on a liquid, ADN-based monopropellant.
Thus, it is an objective of the present invention to eliminate, or at least to suppress hard starts, which are detrimental to the thruster within its specified operating range, and, in addition, to maintain nominal response times for the thruster, since the rise and decay time significantly increases at hard starts when using liquid ammonium dinitramide-based monopropellants.
Other objects and advantages of the present invention will become evident from the following description, examples, and the attached claims.
The terms “rocket engine” and “thruster” will be used interchangeably herein to designate the portion of the inventive liquid propellant rocket engine, into which the propellant is injected, extending downstream to, and including, the nozzle.
The thrust of the inventive rocket engine referred to herein is typically from 0.5 N to a few kN, such as 0.5 to about 3 kN, or 0.5 to 1 kN, and more preferably from 0.5 N to about 500 N.